The present invention generally relates to flow devices used in a fluid system, and more particularly to a flow device that is adapted to restrict flow of a fluid within a system and further capable of responding to an over-pressure condition in the system.
As known in the art, a portion of the air compressed by the fan and/or compressor of a gas turbine engine is typically used for various purposes, including cooling of engine components and, in the case of aircraft engines, use in anti-icing and de-icing systems and cabin pressurization. For illustrative purposes, FIG. 1 schematically represents a high-bypass turbofan engine 10 as including a large fan 12 placed at the front of the engine 10 to compress incoming air 14. The bulk of this compressed air is ducted toward the rear of the engine 10 to increase thrust and reduce the specific fuel consumption of the engine 10, while a smaller portion of the air enters a core engine (gas turbine) 16. The core engine 16 is represented as including a compressor section 18 containing low and high pressure compressor stages that further compress the air, a combustion chamber 20 where fuel is mixed with the compressed air and combusted, and a turbine section 22 where high and low pressure turbines extract energy from the combustion gases to drive, respectively, the high pressure stage of the compressor section 18 and the fan 12 and low pressure stage of the compressor section 18. The fan 12 is surrounded by a nacelle or fan cowling 24 that, in combination with a core cowling 26 surrounding the core engine 16, defines a bypass duct 28 through which the bulk of the compressed fan air flows toward the rear of the engine 10.
FIG. 1 schematically represents bleed air flow that is drawn from the compressor section 18 and delivered through a duct 32 to an annular-shaped cavity 34, sometimes referred to as the D-duct, defined between the inlet lip 36 of the fan cowling 24 and a bulkhead 38 within the fan cowling 24. The hot bleed air can be discharged through a plenum 40 (as represented in FIG. 1) toward the internal surfaces the inlet lip 36, resulting in heating of the lip 36 to remove and/or prevent ice formation. The spent bleed air then exits the D-duct 34 through, for example, one or more vents (not shown). As previously noted, compressed air from the compressor section 18 as well as the fan 12 can be bled for various other purposes and therefore used by various other regions of the engine 10. For example, bleed air from the fan bypass duct 28 is often ducted to the core compartment 30 within the core cowling 26 to cool the engine control and various other components (not shown) located within the compartment 30. As such, the location and configuration of the duct 32 is a nonlimiting example of a duct intended to deliver bleed air within an aircraft engine.
Depending on its end use, air bled from the fan 12 or compressor section 18 may require pressure or flow regulation, such as with a valve or a bleed orifice. For example, the plenum 40 represented in FIG. 1 may be configured as a manifold or one of multiple nozzles (not shown) to serve as a flow restrictor to control the flow rate of bleed air to the D-duct 34. In addition, provisions may be required to ensure that the bleed air does not exceed some pressure limit, for example, the structural capability of the fan cowling 24. For this purpose, various types of pressure relief valves and “blowout” doors have been developed and used. For example, U.S. Pat. No. 3,571,977 discloses a pressure release door to prevent the over-pressurization of a pressurized compartment, such as within a nacelle of an aircraft engine to prevent the failure of skin panels and other major structural components of the nacelle if an over-pressurization event were to occur. While effective for their intended purpose, blowout doors and pressure relief valves add complexity and weight to an aircraft engine, all of which is detrimental to the cost and operation of the aircraft.